JPH10184303A - Stall preventive cascade structure - Google Patents
Stall preventive cascade structureInfo
- Publication number
- JPH10184303A JPH10184303A JP34741396A JP34741396A JPH10184303A JP H10184303 A JPH10184303 A JP H10184303A JP 34741396 A JP34741396 A JP 34741396A JP 34741396 A JP34741396 A JP 34741396A JP H10184303 A JPH10184303 A JP H10184303A
- Authority
- JP
- Japan
- Prior art keywords
- blade
- root
- boundary layer
- back surface
- trailing edge
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Pending
Links
Landscapes
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Description
【0001】[0001]
【発明の属する技術分野】本発明は、軸流機械の翼列の
翼形状に係わり、特に後縁形状をストール防止に有効な
翼形状としたストール防止翼列構造に関する。BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a blade shape of a cascade of an axial flow machine, and more particularly to a stall prevention cascade structure having a trailing edge shape effective for stall prevention.
【0002】[0002]
【従来の技術】軸流機械の翼列では2次流れによる損失
が発生する。図6は軸流機械の翼列に発生する2次流れ
による損失を説明する図である。軸流機械の翼列の隣接
する2つの翼1、2の対向して流体の流路をなす壁面、
すなわち翼1の背面1aと翼2の腹面2bの圧力が異な
るため、翼2の腹面2bに生ずる高圧の境界層7bか
ら、流路底面部3に生ずる境界層7cを通り、翼1の背
面1aに生じる低圧の境界層7aに向かって流れる。こ
の境界層内の流れを2次流れ4と称する。翼1の背面付
け根部1cは主流6の作用により渦流5が発生し易く、
この渦流5により流路底面部3に生ずる境界層7cが翼
1の背面付け根部1cに堆積するため、翼背面における
境界層の剥離(ストール)を起こし、翼性能を低下させ
る。2. Description of the Related Art In a blade cascade of an axial flow machine, a loss occurs due to a secondary flow. FIG. 6 is a diagram illustrating a loss due to a secondary flow generated in a cascade of an axial flow machine. Wall surfaces of two adjacent blades 1 and 2 of the cascade of the axial flow machine forming opposed fluid flow paths;
That is, since the pressure of the back surface 1a of the wing 1 and the pressure of the abdominal surface 2b of the wing 2 are different, the high pressure boundary layer 7b generated on the abdominal surface 2b of the wing 2 passes through the boundary layer 7c generated on the flow path bottom surface portion 3 and passes through the back surface 1a of the wing 1 Flows toward the low-pressure boundary layer 7a. The flow in this boundary layer is referred to as a secondary flow 4. A vortex 5 is easily generated at the base 1c of the back of the wing 1 by the action of the main flow 6,
Since the boundary layer 7c generated on the flow path bottom portion 3 by the vortex 5 accumulates on the rear root portion 1c of the blade 1, the boundary layer is separated (stalled) on the blade rear surface, and the blade performance is reduced.
【0003】特開昭52−67404号公報にはこの境
界層の剥離を防止する方法として、流路底部に翼の背面
側から湾曲した突出部を設け、流路底部を階段状に構成
して翼背面における境界層の剥離を防止する技術が示さ
れている。また、特開昭52−54808号公報には流
路底部に、流路途中の流れの転向する部分以降に1個以
上の障害物を設け、翼背面における境界層の剥離を防止
する技術が示されている。Japanese Patent Application Laid-Open No. 52-67404 discloses a method for preventing the separation of the boundary layer by providing a curved portion from the rear side of the blade at the bottom of the flow path, and forming the bottom of the flow path in a stepwise manner. A technique for preventing separation of a boundary layer at the back of the wing is disclosed. Japanese Patent Application Laid-Open No. 52-54808 discloses a technique in which one or more obstacles are provided at the bottom of the flow path after the turning point of the flow in the middle of the flow path to prevent separation of the boundary layer on the back surface of the blade. Have been.
【0004】[0004]
【発明が解決しようとする課題】両公報に開示された技
術は流路底部に突出部を設けたり、障害物を設けたりす
るため部品数が増大し製作費用が嵩むという問題点があ
った。本発明は、かかる問題点に鑑みてなされたもの
で、翼後縁の形状を変えることにより翼背面の付け根や
先端に生ずる境界層の剥離を防止することを目的とす
る。The technologies disclosed in the two publications have a problem that the number of parts is increased due to the provision of a protruding portion at the bottom of the flow path and the provision of obstacles, and the production cost is increased. The present invention has been made in view of such a problem, and an object of the present invention is to prevent separation of a boundary layer generated at a root or a tip of a back surface of a blade by changing a shape of a trailing edge of the blade.
【0005】[0005]
【課題を解決するための手段】上記目的を達成するた
め、請求項1の発明では、軸流機械の翼列構造におい
て、翼後縁を中央より付け根に向かって拡張してゆく翼
形状とする。In order to achieve the above object, according to the first aspect of the present invention, in the cascade structure of the axial flow machine, the trailing edge of the blade has a blade shape extending from the center toward the root. .
【0006】翼背面の後縁側で翼付け根部は図2(A)
に示すように境界層が剥離し易い。図1は翼後縁を中央
から付け根に向けておよび先端に向けて拡張した形状を
有する翼の翼背面静圧等圧線を示す。翼背面の流速は後
縁に向かうに従い減速するので、静圧は大きくなってゆ
く。静圧の差が翼に働く力となりこの反力Fが流体に働
く。このFの翼付け根方向の成分Fyにより境界層の剥
離を抑えることができる。FIG. 2 (A) shows the wing root at the trailing edge of the back of the wing.
As shown in (1), the boundary layer is easily peeled. FIG. 1 shows a wing back static pressure isobar of a wing having a shape in which the wing trailing edge is expanded from the center toward the root and toward the tip. Since the flow velocity on the back of the blade slows down toward the trailing edge, the static pressure increases. The difference in static pressure becomes a force acting on the wing, and this reaction force F acts on the fluid. The separation of the boundary layer can be suppressed by the component Fy in the blade root direction of F.
【0007】請求項2の発明では、軸流機械の翼列構造
において、翼後縁を中央より先端に向かって拡張してゆ
く翼形状とする。According to the second aspect of the present invention, in the cascade structure of the axial flow machine, the trailing edge of the blade has a blade shape that expands from the center toward the tip.
【0008】翼背面の後縁側で翼先端部も付け根部と同
様に境界層が剥離し易い。翼後縁を中央から先端に向け
て拡張した形状とすることにより、請求項1の発明の付
け根を拡張した場合と同様に流体に翼先端方向の成分の
力が働き、境界層の剥離を抑えることができる。At the trailing edge of the rear surface of the blade, the boundary layer is easily peeled off at the blade tip similarly to the root. By forming the trailing edge of the blade from the center toward the tip, the force of the component in the blade tip direction acts on the fluid in the same manner as in the case where the root of the invention of claim 1 is expanded, thereby suppressing separation of the boundary layer. be able to.
【0009】請求項3の発明では、軸流機械の翼列構造
において、翼後縁を中央より付け根および先端に向かっ
てそれぞれ拡張してゆく翼形状とする。According to a third aspect of the present invention, in the cascade structure of the axial-flow machine, the trailing edge of the blade has a blade shape that expands from the center toward the root and the tip.
【0010】翼後縁を中央より付け根および先端に向か
ってそれぞれ拡張してゆく翼形状とすることにより、静
圧により翼に働く力の反力が流体に働き、この反力の翼
付け根方向の分力と翼先端方向の分力により、翼背面の
付け根部および先端部に発生する剥離を押さえることが
できる。[0010] By forming the trailing edge of the blade into a blade shape that expands from the center toward the root and the tip, the reaction force of the force acting on the blade by the static pressure acts on the fluid, and the reaction force in the direction of the root of the blade is generated. By the component force and the component force in the direction of the blade tip, separation occurring at the root and tip of the blade back surface can be suppressed.
【0011】[0011]
【発明の実施の形態】以下、本発明の実施の形態につい
て図面を参照して説明する。図1は本発明の実施形態を
示す翼側面図で、(A)は翼背面の静圧等圧線とこの静
圧によって流体に生じる反力を示し、(B)は翼後縁拡
張部の詳細図である。翼はターボ圧縮機の翼列を構成す
るもので、動翼、静翼いずれにも適用される。動翼の場
合、翼付け根部はハブに取付けられ、先端はシュラウド
に取付けられる。静翼の場合、翼付け根部はハブに取付
けられ、先端はケーシングに取付けられる。なお、隣接
する2枚の翼と、翼付け根を取付けるハブ、翼先端を取
付けるシュラウドまたはケーシングは流路を形成するの
で、翼付け根側をインナ側流路、翼先端側をアウタ側流
路と称する。Embodiments of the present invention will be described below with reference to the drawings. 1A and 1B are side views of a blade showing an embodiment of the present invention, in which FIG. 1A shows a static pressure isobar on the back surface of the blade and a reaction force generated in the fluid by the static pressure, and FIG. It is. The blades constitute a cascade of a turbo compressor, and are applied to both moving blades and stationary blades. In the case of a bucket, the root of the blade is attached to the hub and the tip is attached to the shroud. In the case of the vane, the root of the wing is attached to the hub, and the tip is attached to the casing. Since two adjacent blades, a hub for mounting the root of the blade, and a shroud or casing for mounting the blade tip form a flow channel, the root of the blade is referred to as an inner flow channel, and the tip of the blade is referred to as an outer flow channel. .
【0012】翼背面では流速は前縁より後縁に行くに従
い減速してゆくので、静圧は前縁より後縁に行くに従い
大きくなってゆく。さらに翼後縁が中央より付け根にか
けて湾曲して広がっており、同様に翼後縁が中央より先
端にかけて湾曲して広がっているので静圧等圧線も湾曲
している。翼背面に加わる静圧の差は力となり、この力
の反力Fが流体に働く。この反力Fは静圧等圧線に直角
に働くため、翼弦長方向前縁より中央部までは流れと反
対方向に働く。中央部より後縁で付け根側および先端側
では湾曲に直角に力Fが働く。力Fの翼付け根方向また
は翼先端方向の分力Fyは境界層の剥離を押さえるよう
に働く。At the back of the wing, the flow velocity decreases as it goes from the leading edge to the trailing edge, so that the static pressure increases as it goes from the leading edge to the trailing edge. Further, the trailing edge of the blade is curved and spread from the center to the root, and similarly, the trailing edge of the blade is curved and spread from the center to the tip, so that the static pressure isobar is also curved. The difference in the static pressure applied to the blade back surface becomes a force, and the reaction force F of this force acts on the fluid. Since this reaction force F acts at right angles to the static pressure isobar, it acts in the direction opposite to the flow from the leading edge in the chord length direction to the center. A force F acts at right angles to the curvature on the root side and the tip side on the trailing edge from the center. The component force Fy of the force F in the blade root direction or the blade tip direction acts to suppress separation of the boundary layer.
【0013】反力Fの大きさは付け根部や先端部の広が
りの大きさと湾曲形状を考慮して静圧等圧線を算出して
求めることができる。The magnitude of the reaction force F can be obtained by calculating a static pressure isobar in consideration of the extent of the root and the tip and the curved shape.
【0014】図2の(A)は従来の翼形の境界層剥離状
態を示し、(B)は本発明の境界層の状態を示す。後縁
は付け根部のみ広げた場合を示すが、先端を広げた場合
も同様な効果が得られる。FIG. 2A shows a state of separation of a conventional airfoil boundary layer, and FIG. 2B shows a state of the boundary layer of the present invention. The trailing edge shows a case where only the base portion is widened, but the same effect can be obtained when the tip is widened.
【0015】図3は図2に示した従来の翼型と、本発明
の付け根部を広げた場合の圧力損失係数を示す。圧力損
失係数は境界層の剥離が発生すると大きくなるので、こ
の剥離を抑制する機能を示す指標として用いることがで
きる。(A)の横軸の%スパンとは(B)に示すように
翼の高さを表し、インナ側通路はスパン%を0%とし、
アウタ側通路は100%とする。本翼形状は付け根側の
み広げているので0%に近づくにつれて、圧力損失係数
は、従来技術では境界層が剥離するため大きくなるが、
本発明では境界層の剥離を抑えているので、大きくなっ
ていない。FIG. 3 shows the conventional airfoil shown in FIG. 2 and the pressure loss coefficient when the base of the present invention is expanded. Since the pressure loss coefficient increases when the boundary layer is separated, it can be used as an index indicating a function of suppressing the separation. The (%) of the horizontal axis of (A) represents the height of the blade as shown in (B), and the inner side passage has a span% of 0%.
The outer passage is 100%. Since the main wing shape is expanded only on the root side, as it approaches 0%, the pressure loss coefficient increases in the prior art because the boundary layer is separated,
In the present invention, since the separation of the boundary layer is suppressed, the size does not increase.
【0016】図4は翼後縁の付け根側を広げない従来の
翼形状の静圧等圧線を示し、図5は翼後縁の付け根側を
広げた本発明の翼形状の静圧等圧線を示す。図4におい
て、翼付け根側のa部は静圧が極めて大きくなっている
範囲を示し、b部は境界層の剥離が発生している範囲を
示す。また翼上端のc部にもb部よりは小さな境界層の
剥離が発生していることを示す。図5において、図示の
ように翼後縁の付け根側を広げると、a部、b部の静圧
等圧線の乱れは整形され、b部に発生していた境界層の
剥離はほぼ消滅している。なお、c部については後縁の
形状は図4と同じであるので同様の小さな境界層剥離が
発生している。FIG. 4 shows a static pressure isobar of a conventional wing shape in which the root side of the wing trailing edge is not widened, and FIG. 5 shows a static pressure isobar of the wing shape of the present invention in which the root side of the wing trailing edge is widened. In FIG. 4, the part a on the wing root side shows a range where the static pressure is extremely large, and the part b shows a range where the boundary layer is separated. Also, it is shown that the boundary layer peeling smaller than the part b has also occurred in the part c at the upper end of the blade. In FIG. 5, when the root side of the trailing edge of the blade is widened as shown in the drawing, the disturbance of the static pressure isobars at the portions a and b is shaped, and the separation of the boundary layer generated at the portion b is almost eliminated. . Since the shape of the trailing edge of the portion c is the same as that of FIG. 4, similar small boundary layer peeling has occurred.
【0017】このように翼後縁を広げることにより、翼
背面の翼付け根部や上端部に発生する境界層剥離を抑制
することができる。しかし、拡張部があまり大きくなる
と摩擦抵抗が増大し翼効率が減少する等の問題が発生す
るため、この後縁の広げ具合は、まず後縁の形状を設定
し、この形状の翼背面の静圧等圧線を計算や必要に応じ
実験も行って定め、図5に示すように乱れや極度に集中
した分布のない状態になるようにトライアンドエラーで
翼形状を決定する。なお、図2〜図5の説明では翼付け
根部のみ後縁を拡張した場合を示したが、翼上端部を拡
張することにより翼上端部に発生する境界層の剥離を抑
制することができる。By widening the trailing edge of the blade in this manner, boundary layer separation occurring at the root and upper end of the blade on the back surface of the blade can be suppressed. However, if the expansion portion becomes too large, problems such as an increase in frictional resistance and a decrease in blade efficiency occur.Therefore, in order to widen the trailing edge, first set the shape of the trailing edge, and then set the static shape of the rear surface of the blade in this shape. The pressure isobars are determined by calculation and experiments as required, and the blade shape is determined by trial and error so that there is no turbulence or extremely concentrated distribution as shown in FIG. In addition, in the description of FIGS. 2 to 5, the case where the trailing edge is expanded only at the root of the wing is shown.
【0018】[0018]
【発明の効果】以上の説明より明らかなように、本発明
は、翼後縁の翼付け根部や上端部を拡張することによ
り、それらの場所に発生する境界層の剥離を抑制するこ
とができる。これにより圧縮機の効率向上およびサージ
ングに対する余裕(マージン)が増大する。また翼の形
状を変更するだけなので、従来の境界層剥離防止構造に
比べ、部品の増加もなく製作が容易になる。As is apparent from the above description, the present invention can suppress the separation of the boundary layer occurring at those places by expanding the wing roots and the upper ends of the wing trailing edges. . This increases the efficiency of the compressor and increases the margin for surging. Further, since only the shape of the wing is changed, the manufacture becomes easier without increasing the number of parts as compared with the conventional boundary layer separation preventing structure.
【図1】本発明の実施形態を示す翼背面の側面図であ
る。FIG. 1 is a side view of a back surface of a wing showing an embodiment of the present invention.
【図2】本発明の境界層剥離防止効果を説明する図であ
る。FIG. 2 is a diagram illustrating an effect of preventing boundary layer peeling of the present invention.
【図3】圧力損失係数により本発明の効果を説明する図
である。FIG. 3 is a diagram illustrating an effect of the present invention by a pressure loss coefficient.
【図4】従来の翼形状の静圧等圧線を示す図である。FIG. 4 is a diagram showing a static pressure isobar of a conventional wing shape.
【図5】本発明の翼形状の静圧等圧線を示す図である。FIG. 5 is a diagram showing a wing-shaped static pressure isobar of the present invention.
【図6】翼列に発生する2次流れを説明する図である。FIG. 6 is a diagram illustrating a secondary flow generated in a cascade.
Claims (3)
中央より付け根に向かって拡張してゆく翼形状としたこ
とを特徴とするストール防止翼列構造。1. A stall prevention cascade structure in an axial flow machine cascade structure, wherein a trailing edge of the blade is formed in a blade shape extending from a center toward a root.
中央より先端に向かって拡張してゆく翼形状としたこと
を特徴とするストール防止翼列構造。2. A stall prevention cascade structure in an axial flow machine cascade structure, wherein a trailing edge of the blade is formed in a blade shape extending from a center toward a tip.
中央より付け根および先端に向かってそれぞれ拡張して
ゆく翼形状としたことを特徴とするストール防止翼列構
造。3. A stall prevention cascade structure in an axial flow machine cascade structure, wherein a trailing edge of the blade is formed in a blade shape extending from a center toward a root and a tip, respectively.
Priority Applications (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
JP34741396A JPH10184303A (en) | 1996-12-26 | 1996-12-26 | Stall preventive cascade structure |
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
JP34741396A JPH10184303A (en) | 1996-12-26 | 1996-12-26 | Stall preventive cascade structure |
Publications (1)
Publication Number | Publication Date |
---|---|
JPH10184303A true JPH10184303A (en) | 1998-07-14 |
Family
ID=18390060
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
JP34741396A Pending JPH10184303A (en) | 1996-12-26 | 1996-12-26 | Stall preventive cascade structure |
Country Status (1)
Country | Link |
---|---|
JP (1) | JPH10184303A (en) |
Cited By (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
WO2007105380A1 (en) | 2006-03-14 | 2007-09-20 | Mitsubishi Heavy Industries, Ltd. | Blade for axial-flow fluid machine |
JP2011099452A (en) * | 2011-02-21 | 2011-05-19 | Mitsubishi Heavy Ind Ltd | Blade for axial flow fluid machine |
-
1996
- 1996-12-26 JP JP34741396A patent/JPH10184303A/en active Pending
Cited By (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
WO2007105380A1 (en) | 2006-03-14 | 2007-09-20 | Mitsubishi Heavy Industries, Ltd. | Blade for axial-flow fluid machine |
US8100658B2 (en) | 2006-03-14 | 2012-01-24 | Mitsubishi Heavy Industries, Ltd. | Axial-flow fluid machine blade |
JP2011099452A (en) * | 2011-02-21 | 2011-05-19 | Mitsubishi Heavy Ind Ltd | Blade for axial flow fluid machine |
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